In gas turbine jet engines, housing components must be cooled to bring the housing components to a tolerable temperature for the housing, since the housing is subjected to high temperatures by the combustion processes in the combustion chamber and the resulting hot combustion gases. Moreover, the cooling of the housing components is also used to adjust, for example, the size of the clearances between the turbine moving blades of the low-pressure turbine and a sealing structure situated on a housing structure, the so-called “outer air seal.” The adjustment of this clearance between the turbine moving blades and the sealing structure on the housing influences the efficiency of the low-pressure turbine, and it varies depending on the load on the gas turbine jet engine, since the temperatures and the centrifugal forces occurring in the turbine blades, and thus the corresponding dimensions, may change under different loads on the engine. For this reason, it is already known to use a so-called active clearance control (ACC), in which, in a gas turbine jet engine having a multi-shaft dual-flow design, cooling air is taken from a bypass flow or secondary flow, which is generated by a front fan, and directed through corresponding lines and openings in the housing structure, between the bypass flow and main flow onto the inner wall of the housing structure which is situated in the main flow in the area of the low-pressure turbine.
DE 35 409 43 A1 shows a gas turbine jet engine illustrated in FIG. 1, which includes, in sequence from left to right, a front fan 1, a high pressure compressor 2, an annular combustion chamber 3, and a high pressure compressor drive turbine 4, downstream from which a low-pressure turbine 5 is aero-thermodynamically connected for the purpose of driving front fan 1.
Front fan 1 is coupled with low-pressure turbine 5 via a shared inner rotor system 6. In the high pressure or gas generator part, high pressure compressor 2 and associated compressor drive turbine 4 are coupled with each other via a shared rotor system 7. Rotor system 7 coaxially surrounds part of rotor system 6. The principle portion of the air flow supplied by front fan 1 (bypass or secondary air flow S) is fed into secondary channel 8 of the engine for the purpose of generating propulsion thrust; a remaining portion S′ of the air flow supplied by front fan 1 reaches high pressure compressor 2 of the gas generator. The hot gas flow escaping from low-pressure turbine 5 is also used to generate propulsion thrust.
In an engine of this type, principle turbine components must normally be cooled for the purpose of managing the hot gas temperatures. For example, it would be possible to cool the inlet stationary blades of high pressure turbine 4 and also, for example, the moving blades of high pressure turbine 4 as well as, if necessary, for example, the stationary blades of the second stage of high pressure turbine 4. The compressor air used for the aforementioned cooling situations of high pressure turbine 4 may be taken from one or multiple suitable locations of high pressure compressor 2 and made available for the appropriate application, e.g., via the applicable inner rotor system. It is known to apply cooling air to relevant ring-shaped housing structures 9 and 10 of high pressure and low-pressure turbines 4 and 5 via comparatively costly and complicated pipe distribution systems, this cooling air being taken from the secondary channel of the engine.
In the engine in FIG. 1, a secondary air portion removed from the bypass flow is furthermore taken from the secondary flow via openings 11 situated in wall 12 of secondary channel 8 in the immediate vicinity of the relevant turbine housing structure (e.g., low-pressure turbine 5 in this case) for the purpose of cooling the turbine components and optimizing the blade clearance, and this secondary air portion is blown out against turbine housing structure 10 by way of an impact cooling (arrow F).